This invention relates to ceramic coatings on substrates, and, more particularly, to aircraft gas turbine components protected by such coatings.
In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine by impingement against an airfoil section of the turbine blades and vanes, which turns the shaft and provides power to the compressor. In a more complex version of the gas turbine engine, the compressor and a high pressure turbine are mounted on one shaft, and a bypass fan and low pressure turbine are mounted on a separate shaft. In any event, the hot exhaust gases flow from the back of the engine, driving it and the aircraft forwardly.
The hotter the combustion and exhaust gases, the more efficient is the operation of the jet engine. There is therefore an incentive to raise the combustion and exhaust gas temperatures. The maximum temperature of the combustion gases is normally limited by the materials used to fabricate the hot-section components of the engine. These components include the turbine vanes and turbine blades of the gas turbine, upon whose airfoil sections the hot combustion gases directly impinge. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at temperatures of up to about 1800-2100xc2x0 F. These components are also subject to damage by oxidation and corrosive agents, as well as impact damage and erosion by particles entrained in the combustion gas stream.
Many approaches have been used to increase the operating temperature limit and service lives of the turbine blades and vanes to their current levels, while achieving acceptable oxidation, corrosion, erosion, and impact resistance. The composition and processing of the base materials themselves have been improved. Cooling techniques are used, as for example providing the component with internal cooling passages through which cooling air is passed.
In another approach used to protect the hot-section components, the surfaces of the turbine blades and vanes are coated with thermal barrier coating systems. The thermal barrier coating systems typically include a bond coat that contacts the substrate, and a ceramic thermal barrier coating (TBC) layer overlying the bond coat. The bond coat protects the articles against the oxidative and corrosive effects of the combustion gas. The ceramic layer provides thermal insulation and erosion/impact damage resistance. The turbine blades and turbine vanes are thereby able to run cooler and are more resistant to environmental attack in the presence of the thermal barrier coating systems.
Although the thermal barrier coating approach is operable, there is opportunity for improvement. It would be desirable to improve the thermal insulation properties of the ceramic thermal barrier coating, as well as to increase its resistance to impact damage. The present invention fulfills this need, and further provides related advantages.
The present invention provides a structure in which a substrate is protected by an overlying ceramic layer. The substrate may be a component of a gas turbine engine such as a turbine blade or turbine vane. The ceramic layer has improved insulation properties as compared with prior ceramic layers, as well as improved resistance to impact damage. The ceramic layer of the invention is compatible with the use of bond coats, and in some circumstances dispenses with the need for a bond coat.
A method of preparing a structure protected by a ceramic coating comprises the steps of providing a substrate having a surface, and depositing a layer comprising a precursor material onto the surface of the substrate. The precursor material comprises a sacrificial ceramic. The method further includes furnishing a reactive gas, the reactive gas being reactive with the sacrificial ceramic to produce a protective ceramic different from the sacrificial ceramic, and thereafter contacting the reactive gas to the layer comprising the precursor material to produce a protective ceramic layer.
Preferably, the sacrificial ceramic is silica, and the reactive gas comprises an aluminum-containing gas such as aluminum chloride gas. The resulting protective ceramic comprises alumina. Desirably, the protective ceramic layer also comprises at least about 20 percent by volume of intraceramic space, which may be empty and porous, or filled partially with metal.
The present approach is compatible with further treatments, such as the formation of a bond coat layer on the surface of the substrate. A layer of an additional material, such as a ceramic sealing layer, may overlie the protective ceramic layer.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings, which illustrate, by way of example, the principles of the invention. The scope of the invention is not, however, limited to this preferred embodiment.